The airfoils of gas turbines are typically made of nickel or cobalt based superalloys which show high resistance against the hot and corrosive combustion gases present in gas turbine. However, although such superalloys have considerably high corrosion and oxidation resistance, the high temperatures of the combustion gases in gas turbines require measures to improve corrosion and/or oxidation resistance further. Therefore, airfoils of gas turbine blades and vanes are typically at least partially coated with a thermal barrier coating system to prolong the resistance against the hot and corrosive environment. In addition, airfoil bodies are typically hollow so as to allow a cooling fluid, typically bleed air from the compressor, to flow through the airfoil. Cooling holes present in the walls of the airfoil bodies allow a certain amount of cooling air to exit the internal passages so as to form a cooling film over the airfoil surface which further protects the superalloy material and the coating applied thereon from the hot and corrosive environment. In particular, cooling holes are present at the trailing edges of the airfoils as it is shown in U.S. Pat. No. 6,077,036, U.S. Pat. No. 6,126,400, US 2009/0194356 A1 and WO 98/10174, for example.
Trailing edge losses are a significant fraction of the over all losses of a turbo machinery blading. In particular, thick trailing edges result in higher losses. For this reason, cooled airfoils with a cutback design at the trailing edge have been developed. This design is realised by taking away material on the pressure side of the airfoil from the trailing edge up to several millimeters towards the leading edge. This measure provides very thin trailing edges which can provide big improvements on the blading efficiency. An airfoil with a cutback design and a thermal barrier coating is, for example, disclosed in WO 98/10174 A1. However, the beneficial effect on the efficiency can only be achieved if the thickness of the trailing edge is rather small. On the other hand, for a blade with thermal barrier coating the combined thickness of the cast airfoil body wall and the applied thermal barrier coating system exceeds the optimum thickness of the design. In addition, as the flow velocity of the gas is the greatest at the trailing edge of the airfoil a thermal barrier coating applied to the trailing edge is prone to high levels of erosion.
It is known to selectively provide a thermal barrier coating system to the airfoil, in particular such that the trailing edge of an airfoil and adjacent regions of an airfoil remain uncoated. Selective coatings are, for example, described in U.S. Pat. No. 6,126,400, U.S. Pat. No. 6,077,036 and, with respect to the coating method, in US 2009/0104356 A1.
However, in U.S. Pat. No. 6,077,036 the pressure side of the airfoil is completely uncoated which means that areas which would not suffer from a higher combined thickness of the cast airfoil body and the coating applied thereon remain unprotected against the temperature the hot combustion gas.
WO 2008/043340 A1 and US 2010/0014962 A1 describe a turbine airfoil with a thermal barrier coating the thickness of which varies over the airfoil surface. Starting from the flow inlet edge, the layer thickness of the thermal barrier coating on the pressure side decreases continuously in the direction of a flow outlet edge, wherein no thermal barrier coating is preferably applied to the pressure side directly adjacent to the flow outlet edge so that in a section of the pressure side, which as a rule is provided with cooling air exits, the layer thickness of the thermal barrier coating is approximately zero. Part of the pressure side close to the cutback or air gap between the pressure side and the suction side is left uncoated.
In U.S. Pat. No. 6,126,400 the thermal barrier coating only covers about half of the airfoil, as seen from the leading edge towards the trailing edge.
In WO 99/48837 a ceramic composition for insulating components, made of ceramic matrix composites, of gas turbines is provided.
EP 1 544 414 A1 discloses an inboard cooled nozzle doublet, wherein a doublet of hollow vanes is integrally joined between two bands of a turbine nozzle. The vanes comprise rows of trailing edge outlets.
In U.S. Pat. No. 4,121,894 a refurbished turbine vane or blade is disclosed. The refurbished turbine vane or blade comprises an overlay metal which has been added to the vane surfaces by a plasma spray process and thereafter refinished to conform to the original contours as specified for new vanes. The overlay metal can be applied to build up a thickness of as much as 30 to 40 thousands of an Inch, and can be feathered as the overlay approaches the trailing edge of the vane. This means, that the area around the trailing edge is not covered by the overlay metal.
The trailing edge of an aerofoil requires being as thin as possible due to the considerable aerodynamic losses incurred. On a cooled vane, the target thickness for the trailing edge must include two cast wall thicknesses, an air gap and two thermal barrier coating thicknesses. Due to a minimum casting thickness, the sum of all the thicknesses exceeds the overall target. Previously, a similar part has been left uncoated, hence being subject to higher oxidation.